1. Field of the Invention
The present invention relates generally to air cooled turbine airfoils, and more specifically to a process for modeling of specific detailed features including turbulators, impingement holes and film holes in a turbine airfoil.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
A gas turbine engine includes a turbine section with at least one stage of stator vanes and rotor blades that are exposed to a high temperature gas flow from the combustor for the extraction of mechanical energy used to drive the compressor of the engine, or in the case of an industrial gas turbine engine (IGT) to also drive an electric generator. It is well known in the art of gas turbine engines that the efficiency of the engine can be increased by passing a higher temperature gas flow into the turbine. However, a limit to the highest temperature is imposed by the material properties of the first stage airfoils and to the amount of internal cooling provided to these airfoils. Later stages of airfoils also may require cooling but are not as limiting as to the turbine inlet temperature as are the first stage airfoils because the first stage is exposed to the highest temperature.
A cooled turbine blade design for a heavy-duty gas turbine has major specification targets to be achieved, which include performance, emissions, lifetime, and cost and development time. The front stages of the gas turbine are exposed to the most severe conditions in terms of loading, specifically the thermal loads, requiring materials such as nickel based single crystal alloys, which result in associated costs in the manufacturing supply chain. The consequences of working closer to an achievable technology limit in general results in engineering design of new components being complex, engineering development time raises and constraints become more severe, overall investment risks increase, individual process steps become more dependent on each other and interfaces get more complex, and the available response time to field issues are reduced. Desired inlet temperatures are about 2150 K for an aircraft engine and 1850 K for an IGT, which far exceed the allowable material temperatures of 1320 to 1475 K for super alloys without thermal barrier coatings (TBC) and 1475 to 1625 K with a TBC. But, effective in addition to efficient cooling are needed because the slightest over temperature in any part of the turbine material such as a hot spot in a blade or vane from insufficient cooling will lead to failure of that part.
Lifting of cooled turbine blades and vanes is also an important design factor in a gas turbine engine. These airfoils are exposed to cycles of high operating temperature with high stress levels due to rotation and/or pressure forces followed by low temperatures and forces when the engine is shut down. Lifting of cooled turbine blades requires three-dimensional finite element models to be available which represent the part detail in a physically reduced complexity which can be handled with available methods and computing power. It takes development engineers involved (mainly aerodynamic design, mechanical design, cooling and mechanical integrity) several months to get an iterated feasible design solution, where, what is often called an “optimum solution”, has been allowed only limited iterations on the overall three-dimensional level involving all interfaces in between disciplines.
The design goal for a turbine blade (or vane) is to reduce the temperature of the blade surface to within the material capability using a minimum amount of cooling flow. The internal geometry is designed to sufficiently cool the blade allowing it to efficiently operate at very high temperatures. Cooling fluid extracted from the compressor is circulated through internal serpentine passages that internally cool the blade by a combination of convection and impingement cooling. The design intent in the internal passages is to maximize the heat transfer coefficient. Flow is extracted from the internal cooling circuit and ejected through numerous small holes located near the leading edge, as well as along the pressure side and suction side of the external surface for film cooling purposes. The turbine blade geometry and associated fluid dynamics and thermodynamics is complex, three-dimensional and unsteady. The internal flow-fields alone can be quite challenging to study computationally due to the internal serpentine cooling circuit. The external flow-field is challenging to model computationally due in part to the scales of the film cooling holes in relation to the overall geometry.
Traditionally, the heat transfer problem has been studied by decoupling the external flow, internal flow and film cooling holes. By this, we mean that the modeling of the external flow, internal flow and film cooling holes is done separately. Thermal modeling of a modern convective/film cooled gas turbine airfoil has traditionally been distributed over three domains. They include a flow path fluid domain (calculated with correlation or computational fluid dynamics or CFD), an internal cooling fluid domain (calculated with correlation or CFD), and an airfoil solid domain (calculated with finite element analysis or FEA). The three domains must communicate together in order to achieve a balance of mass flow and thermal energy. This communication between the three domains has typically been achieved through script driven transfer of thermal/flow boundary condition information, which is a time consuming and error prone process.
Recent modeling of cooled turbine airfoils include the use of conjugate heat transfer in which not only is the heat transfer in the turbine material accounted for but also the coupling between the internal and the film cooling of the turbine material. Conjugate heat transfer modeling simultaneously solves all three domains in a single coupled model, potentially saving time and virtually eliminating heat balance errors between the three domains. The metal temperature and gas pressure information may be passed directly to the FEA mechanical model. However, prior art conjugate simulations for entire internal and film cooled turbine blades are few due to the high computational cost.
A majority of the cooling air (and therefore the design effort) in modern turbines is used by the first and second stage airfoils for cooling. These cooling designs tend to have a high convective efficiency internal cooling schemes that exhaust cooling air in the form of film cooling. A modern cooled industrial power turbine airfoil may have multiple serpentine passages feeding about 1000 film holes. To further add to the complexity, internal cooling effectiveness is augmented by around 100 turbulators or around 2000 impingement holes. All features are important to the thermal model and cannot be neglected without compromising physical accuracy. However, conjugate CFD analysis of such a configuration is not practical due to the meshing effort and the computational demands.
The primary difficulty associated with conjugate analysis of the aforementioned modern airfoil is a result of miss-matched length scales. For example, an industrial gas turbine airfoil may have a passage height of about 200 mm which is a manageable solution by industry standards but approaching the upper limit of practicality for use in an iterative design environment. This same 200 mm airfoil may have around 1000 plus film holes with a diameter of around 1.0 mm. It is known that a maximum cell size of 1/10th to 1/50th of the hole diameter is required to capture film hole physics for accurate prediction of film cooling. To decrease average flow path cell size from 1.0 mm to 0.1 mm increases the mesh size by three orders of magnitude. Accordingly, aggressive mesh size transition (high manpower) must be used to keep computational demands within the limits of practicality. Many attempts have been made to automate the meshing process. However, the geometric complexity of internal cooling circuits has largely prevented efficient automation.